XFOIL Version 6.94 Calculated polar for: YS-930 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3607 0.01162 0.00700 -0.1074 0.8952 0.9818 0.500 0.4116 0.01201 0.00734 -0.1071 0.8886 1.0000 1.000 0.4788 0.01190 0.00722 -0.1095 0.8847 1.0000 1.500 0.5235 0.01229 0.00771 -0.1074 0.8760 1.0000 2.000 0.5938 0.01181 0.00730 -0.1097 0.8686 1.0000 2.500 0.6564 0.01119 0.00680 -0.1100 0.8546 1.0000 3.000 0.7287 0.00980 0.00554 -0.1110 0.8310 1.0000 3.500 0.7790 0.00876 0.00457 -0.1075 0.7850 1.0000 4.000 0.8016 0.00888 0.00387 -0.0987 0.5709 1.0000 4.500 0.7597 0.01474 0.00666 -0.0807 0.0239 1.0000 5.000 0.7960 0.01624 0.00820 -0.0771 0.0096 1.0000 5.500 0.8297 0.01839 0.01051 -0.0728 0.0083 1.0000 6.000 0.8731 0.02128 0.01358 -0.0701 0.0081 1.0000 6.500 0.9404 0.02812 0.02111 -0.0699 0.0097 1.0000 7.000 0.9977 0.03882 0.03286 -0.0655 0.0265 1.0000 7.500 1.0096 0.04612 0.04102 -0.0576 0.0223 1.0000 8.000 1.0221 0.05199 0.04746 -0.0515 0.0200 1.0000 8.500 1.0275 0.05767 0.05343 -0.0464 0.0187 1.0000 9.500 0.9824 0.07561 0.07191 -0.0340 0.0175 1.0000 10.000 0.9439 0.08167 0.07830 -0.0280 0.0174 1.0000 10.500 0.9072 0.09022 0.08709 -0.0265 0.0174 1.0000 11.000 0.8734 0.09989 0.09697 -0.0295 0.0174 1.0000 11.500 0.8486 0.11213 0.10937 -0.0370 0.0174 1.0000 12.000 0.8347 0.12977 0.12706 -0.0492 0.0174 1.0000 12.500 0.8343 0.14182 0.13910 -0.0559 0.0174 1.0000 13.000 0.8330 0.15429 0.15152 -0.0623 0.0172 1.0000 13.500 0.8354 0.16633 0.16353 -0.0699 0.0162 1.0000 14.000 0.8410 0.17611 0.17328 -0.0748 0.0152 1.0000 14.500 0.8475 0.18542 0.18258 -0.0790 0.0142 1.0000 15.000 0.8558 0.19406 0.19121 -0.0829 0.0134 1.0000 15.500 0.8658 0.20205 0.19919 -0.0861 0.0124 1.0000 16.000 0.8789 0.20941 0.20656 -0.0885 0.0119 1.0000